Air seal system with backside abradable layer

ABSTRACT

A seal system for a gas turbine engine includes a ceramic matrix composite (CMC) seal arc segment and a carrier supporting the CMC seal arc segment. The CMC seal arc segment defines radially inner and outer sides and has an abradable layer disposed on the radially outer side. There is a cooling cavity radially between the carrier and the abradable layer. The carrier includes a ridge that projects into a groove in the abradable layer and provides a labyrinth seal that partitions the cooling cavity into sub-cavities.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section, and a turbine section. Air entering thecompressor section is compressed and delivered into the combustionsection where it is mixed with fuel and ignited to generate ahigh-pressure and temperature exhaust gas flow. The high-pressure andtemperature exhaust gas flow expands through the turbine section todrive the compressor and the fan section. The compressor section mayinclude low and high pressure compressors, and the turbine section mayalso include low and high pressure turbines.

The turbine section may include a row of air seals positioned near thetips of the turbine blades. The air seals are in close proximity to theblade tips to provide a minimum clearance distance and thereby reduceleakage of combustion gases around the tips. As the air seals areexposed to the hot exhaust gas flow, cooling air is provided to the airseals for thermal management.

SUMMARY

A seal system for a gas turbine engine according to an example of thepresent disclosure includes a ceramic matrix composite (CMC) seal arcsegment that defines radially inner and outer sides, and has anabradable layer disposed on the radially outer side. A carrier supportsthe CMC seal arc segment, and there is a cooling cavity radially betweenthe carrier and the abradable layer. The carrier includes a ridge thatprojects into a groove in the abradable layer and thereby provides alabyrinth seal that partitions the cooling cavity into sub-cavities.

In a further embodiment of any of the foregoing embodiments, the ridgehas a knife edge.

In a further embodiment of any of the foregoing embodiments, theabradable layer is selected from the group consisting of pure silicon,yttria stabilized zirconia, gadolinia stabilized zirconia, andcombinations thereof.

In a further embodiment of any of the foregoing embodiments, the carrierincludes radially inner and outer carrier sides. The radially innercarrier side bounds the sub-cavities. The carrier includes a firstsupply passage that opens at the radially outer carrier side and at theradially inner carrier side to one of the sub-cavities, and the carrierincludes a second supply passage that opens at the radially outercarrier side and at the radially inner carrier side to another,different one of the sub-cavities.

In a further embodiment of any of the foregoing embodiments, the outerside of the CMC arc segment has a pair of hooks that define a dovetailslot there between. The carrier has a dovetail disposed in the dovetailslot, and the first supply passage and the second supply passage arethrough the dovetail.

In a further embodiment of any of the foregoing embodiments, the CMCseal arc segment defines forward and aft sides, one of the hooks of thepair of hooks is elongated along the forward side, and the other of thehooks of the pair of hooks is elongated along the aft side.

In a further embodiment of any of the foregoing embodiments, thesub-cavities are pressurized at different pressures.

In a further embodiment of any of the foregoing embodiments, thecarrier, including the ridge, is selected from the group consisting of aNi-based alloy and a cobalt-based alloy.

A gas turbine engine according to an example of the present disclosureincludes a compressor section, a combustor in fluid communication withthe compressor section, and a turbine section in fluid communicationwith the combustor. The turbine section has a seal system disposed abouta central axis of the gas turbine engine. The seal system includes aceramic matrix composite (CMC) seal arc segment that defines radiallyinner and outer sides and has an abradable layer disposed on theradially outer side. A carrier supports the CMC seal arc segment, andthere is a cooling cavity radially between the carrier and the abradablelayer. The carrier includes a ridge that projects into a groove in theabradable layer and thereby provides a labyrinth seal that partitionsthe cooling cavity into sub-cavities

In a further embodiment of any of the foregoing embodiments, the carrierincludes radially inner and outer carrier sides. The radially innercarrier side bounds the sub-cavities. The carrier includes a firstsupply passage that opens at the radially outer carrier side and at theradially inner carrier side to one of the sub-cavities. The carrierincludes a second supply passage that opens at the radially outercarrier side and at the radially inner carrier side to another,different one of the sub-cavities. The outer side of the CMC arc segmenthas a pair of hooks that define a dovetail slot there between. Thecarrier has a dovetail disposed in the dovetail slot, and the firstsupply passage and the second supply passage are through the dovetail.

In a further embodiment of any of the foregoing embodiments, the ridgehas a knife edge.

In a further embodiment of any of the foregoing embodiments, thesub-cavities are pressurized at different pressures.

In a further embodiment of any of the foregoing embodiments, theabradable layer is selected from the group consisting of pure silicon,yttria stabilized zirconia, gadolinia stabilized zirconia, andcombinations thereof, and the carrier, including the ridge, is selectedfrom the group consisting of a Ni-based alloy and a cobalt-based alloy.

A method according to an example of the present disclosure includesproviding a ceramic matrix composite (CMC) seal arc segment that hasradially inner and outer sides, attaching an abradable layer on theradially outer side of the CMC seal arc segment, providing a carrier tosupport the CMC seal arc segment, the carrier including a ridge, andsliding the CMC seal arc segment relative to the carrier such thatduring the sliding the ridge cuts a groove into the abradable layer. Theridge remains disposed in the groove to thereby provide a labyrinth sealthat partitions the cooling cavity between the carrier and the CMC sealarc segment into sub-cavities.

In a further embodiment of any of the foregoing embodiments, the outerside of the CMC arc segment has a pair of hooks that define a dovetailslot there between. The carrier has a dovetail, and the sliding includessliding the dovetail into the dove slot.

In a further embodiment of any of the foregoing embodiments, the ridgehas a knife edge.

In a further embodiment of any of the foregoing embodiments, theabradable layer is selected from the group consisting of pure silicon,yttria stabilized zirconia, gadolinia stabilized zirconia, andcombinations thereof, and the carrier, including the ridge, is selectedfrom the group consisting of a Ni-based alloy and a cobalt-based alloy.

The present disclosure may include any one or more of the individualfeatures disclosed above and/or below alone or in any combinationthereof.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of the present disclosure willbecome apparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

FIG. 1 illustrates a gas turbine engine.

FIG. 2 illustrates a portion of the turbine section of the engine.

FIG. 3A illustrates an isometric view of a seal arc segment.

FIG. 3B illustrates a sectioned view of the seal arc segment.

FIG. 4A illustrates an isometric view of a carrier that supports theseal arc segment.

FIG. 4B illustrates a radial view of the carrier.

In this disclosure, like reference numerals designate like elementswhere appropriate and reference numerals with the addition ofone-hundred or multiples thereof designate modified elements that areunderstood to incorporate the same features and benefits of thecorresponding elements.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flow path B in a bypass duct defined within a housing15 such as a fan case or nacelle, and also drives air along a core flowpath C for compression and communication into the combustor section 26then expansion through the turbine section 28. Although depicted as atwo-spool turbofan gas turbine engine in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with two-spool turbofans as the teachings may beapplied to other types of turbine engines including three-spoolarchitectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects, a first (or low) pressure compressor 44 and a first (orlow) pressure turbine 46. The inner shaft 40 is connected to the fan 42through a speed change mechanism, which in exemplary gas turbine engine20 is illustrated as a geared architecture 48 to drive a fan 42 at alower speed than the low speed spool 30. The high speed spool 32includes an outer shaft 50 that interconnects a second (or high)pressure compressor 52 and a second (or high) pressure turbine 54. Acombustor 56 is arranged in the exemplary gas turbine 20 between thehigh pressure compressor 52 and the high pressure turbine 54. Amid-turbine frame 57 of the engine static structure 36 may be arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 57 further supports bearing systems 38in the turbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded through the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of the low pressure compressor, or aftof the combustor section 26 or even aft of turbine section 28, and fan42 may be positioned forward or aft of the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), andcan be less than or equal to about 18.0, or more narrowly can be lessthan or equal to 16.0. The geared architecture 48 is an epicyclic geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3. The gear reduction ratio maybe less than or equal to 4.0. The low pressure turbine 46 has a pressureratio that is greater than about five. The low pressure turbine pressureratio can be less than or equal to 13.0, or more narrowly less than orequal to 12.0. In one disclosed embodiment, the engine 20 bypass ratiois greater than about ten (10:1), the fan diameter is significantlylarger than that of the low pressure compressor 44, and the low pressureturbine 46 has a pressure ratio that is greater than about five 5:1. Lowpressure turbine 46 pressure ratio is pressure measured prior to aninlet of low pressure turbine 46 as related to the pressure at theoutlet of the low pressure turbine 46 prior to an exhaust nozzle. Thegeared architecture 48 may be an epicycle gear train, such as aplanetary gear system or other gear system, with a gear reduction ratioof greater than about 2.3:1 and less than about 5:1. It should beunderstood, however, that the above parameters are only exemplary of oneembodiment of a geared architecture engine and that the presentinvention is applicable to other gas turbine engines including directdrive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. The engine parameters described above and those in thisparagraph are measured at this condition unless otherwise specified.“Low fan pressure ratio” is the pressure ratio across the fan bladealone, without a Fan Exit Guide Vane (“FEGV”) system. The low fanpressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45, or more narrowly greater than orequal to 1.25. “Low corrected fan tip speed” is the actual fan tip speedin ft/sec divided by an industry standard temperature correction of[(Tram ° R)/(518.7° R)]^(0.5). The “Low corrected fan tip speed” asdisclosed herein according to one non-limiting embodiment is less thanabout 1150.0 ft/second (350.5 meters/second), and can be greater than orequal to 1000.0 ft/second (304.8 meters/second).

FIG. 2 illustrates a portion from the turbine section 28 of the engine20. The turbine section 28 includes a seal system 60 that is situatedradially outwardly of a row of turbine blades (not shown). For example,the seal system 60 may function to provide a minimal clearance distancewith the tips of the blades to facilitate a reduction in leakage ofcombustion gases around the tips. In this example, the seal system 60 isdirectly attached to a turbine case 62, although there may alternativelybe one or more intermediate structures between the seal system and thecase 62.

The seal system 60 includes a row of ceramic matrix composite (CMC) sealarc segments 64 (one shown) disposed around the engine centrallongitudinal axis A (superimposed on the figure) and a carrier 66 thatsupports the segment 64 via attachment to the case 62. Referring also toFIGS. 3A and 3B that show an isolated view of the segment 64 and across-section of the segment 64, respectively, the segment 64 is formedby a main body 68 that defines radially inner and outer sides 68 a/68 b,forward and aft sides 68 c/68 d, and circumferential sides 68 e/68 f.The radially inner side 68 a faces the core gaspath C.

In this example, on the outer side 68 b the segment 64 includes a pairof hooks 70 that slope inwards toward the axial midline of the segment64 and define a dovetail slot 72 there between. The forward one of thehooks 70 is elongated along the forward side 68 c, and the aft one ofthe hooks 70 is elongated along the aft side 68 d. Each of the hooks 70extends the full circumferential distance from one circumferential side68 e to the other circumferential side 68 f.

The CMC material from which the CMC vane arc segments 64 are made of iscomprised of a ceramic reinforcement, which is usually continuousceramic fibers, in a ceramic matrix. Example ceramic matrices aresilicon-containing ceramic, such as but not limited to, a siliconcarbide (SiC) matrix or a silicon nitride (Si3N4) matrix. Exampleceramic reinforcement of the CMC are silicon-containing ceramic fibers,such as but not limited to, silicon carbide (SiC) fiber or siliconnitride (Si3N4) fibers. The CMC may be, but is not limited to, a SiC/SiCceramic matrix composite in which SiC fiber plies are disposed within aSiC matrix. The fiber plies have a fiber architecture, which refers toan ordered arrangement of the fiber tows relative to one another, suchas a 2D woven ply or a 3D structure.

The illustrated geometry of the segment 64 is amenable to ceramicprocessing. For instance, the segment 64 may be fabricated by laying-upfiber plies on a mandrel or tool that replicates the geometry of thedovetail slot 72. The fiber plies may be wrapped around the mandrel toform the hooks 70. The mandrel may then be removed and the green fiberstructure is densified with the ceramic matrix, such as by chemicalvapor infiltration or polymer impregnation and pyrolysis.

Referring also to FIGS. 4A and 4B that show, respectively, an isometricview of the carrier 66 and a radial view (from the axis A lookingradially outwards) of the carrier 66, the carrier 66 includes a mainbody 74 that defines radially inner and outer carrier sides 74 a/74 b,forward and aft carrier sides 74 c/74 d, and circumferential carriersides 74 e/74 f. In this example, the carrier 66 includes hooks 76 onthe radially outer carrier side 74 b that serve to attach the carrier 66to the case 62. The forward and aft carrier sides 74 c/74 d are slopedso as to form a dovetail 78. The dovetail 78 is complementary ingeometry to the dovetail slot 72 so as to be slidably receivable intothe dovetail slot 72 to secure the segment 64 and the carrier 66together. As shown in FIG. 2 , the segment 64 “hangs” on the carrier 66such that there is a cooling cavity 80, i.e., a plenum, defined betweenthe inner carrier side 74 a and the outer side 68 b of the segment 64.The carrier 66 is made of a metallic alloy, such as a nickel-based orcobalt-based superalloy.

To facilitate cooling of the seal system 60, cooling air, such as bleedair from the compressor section 24, is provided to the backside of thecarrier 66 (outer carrier side 74 b). The carrier 66 includes supplypassages 82 that each open to the outer carrier side 74 b and to theinner carrier side 74 a to deliver the cooling air into the coolingcavity 80. At least a portion of the supply passages 82 extend throughthe dovetail 78.

The segment 64 includes an abradable layer 84 disposed on the radiallyouter side 68 b. For example, the abradable layer 84 may be selectedfrom, but is not limited to, pure silicon, yttria stabilized zirconia,gadolinia stabilized zirconia, and combinations thereof. The innercarrier side 74 a of the carrier 66 includes one or more ridges 86 (twoshown). In this example, each ridge 86 is elongated in thecircumferential direction and extends fully from one side 74 e to theother side 74 f of the carrier 66. The ridge 86 has a radial height thatexceeds the radial distance between the inner carrier side 74 a and thesurface of the abradable layer 84 such that the ridge 86 cuts a groove84 a into the abradable layer 84. The ridge 86 is abrasive relative tothe abradable layer 84. That is, when the ridge 86 rubs on the abradablelayer 86, the abradable layer 86 is worn by the ridge 86. The ridge 86is not worn, or is at least worn substantially less than the abradablelayer 84. In this regard, the abradable layer 84 may have a hardness, aporosity, or a combination thereof that render it abradable by the ridge86, which as above may be made of nickel-based or cobalt-basedsuperalloy. The term “cut” or variations thereof that are used in thisdisclosure thus refer to the physical capability of the ridge or ridges86 to abrade the abradable layer 84 to form the groove or grooves 84 a.

Each ridge 86 extends into its corresponding groove 84 a and therebypartitions the cooling cavity 80 into sub-cavities, which in thisexample are indicated at 80 a/80 b/80 c. The engagement of the ridge 86into the groove 84 a also provides a labyrinth seal such that thesub-cavities 80 a/80 b/80 c are substantially flow- andpressure-isolated from each other. In this example, as the ridges 86fully extend from side 74 e to side 74 f of the carrier 66, thesub-cavity 80 a is an axially forward sub-cavity, the sub-cavity 80 c isan axially aft sub-cavity, and the sub-cavity 80 b is an axiallyintermediate sub-cavity. That is, the sub-cavities are in an axiallyserial arrangement. The carrier 66 may alternatively have a single ridge86 to provide two sub-cavities, or additional ridges 86 to provide morethan three sub-cavities.

The compartmentalization of the cavity 80 into the sub-cavities 80 a/80b/80 c by the ridge(s) 86 and groove(s) 84 a enables the cooling air tobe tailored across the seal system 60. For example, the cooling passages82 meter flow of the cooling air. Through selection of the size andnumber of the cooling passages 82, the flow and pressure of the coolingair provided to each of the sub-cavities 80 a/80 b/80 c is manipulatedto pressurize the sub-cavities at different pressures.

In one example, the flow and pressure of the cooling air provided toeach of the sub-cavities 80 a/80 b/80 c is manipulated with respect tothe axially changing pressure profile in the core gaspath C. Along theaxial extent of the segment 64, the core gaspath pressure is highest atregions forward of the blades. The core gaspath pressure is lowest atregions aft of the blades, and the core gaspath pressure is at anintermediate level across the blades (between their forward and leadingedges). The cooling air provided to the seal system is pressurized fromthe compressor section 24. The pressure of the cooling air and pressuresin the core gaspath C, and especially the pressure ratio between these,can causes stresses in the segment 64.

The size and number of the cooling passages 82 that feed the cooling airto each of the sub-cavities 80 a/80 b/80 c is selected such that thesub-cavities 80 a are pressurized at progressively lower pressures,i.e., the sub-cavity 80 a has the highest pressure, the sub-cavity hasthe lowest pressure, and the sub-cavity 80 b has an intermediatepressure. The pressure in each sub-cavity 80 a/80 b/80 c is coordinatedto the core gaspath pressure at the axial location of the sub-cavity 80a/80 b/80 c along the core gaspath C to facilitate reduction in thelocal pressure ratio at that sub-cavity 80 a/80 b/80 c, to therebyfacilitate stress reduction in the segment 64 due to the pressure ratio.In terms of flow-metering, the number and/or size of the coolingpassages 82 feeding the sub-cavity 80 a will be greater than that of thecooling passages 82 feeding the sub-cavity 80 b, which in turn will begreater than that of the cooling passages 82 feeding the sub-cavity 80c. The size and/or number of the cooling passages 82 may also beselected in coordination with cooling requirements of the seal system 60to provide a desired balance of cooling and pressure ratio for stressreduction.

Also disclosed is a method for the seal system 60. The method may beused for an original manufacture of the seal system 60 or as part of arepair or replacement process. The method includes providing the segment64 and attaching the abradable layer 84 on the radially outer side 68 bof the segment 64. For instance, the abradable layer 84 is a coatingthat is deposited, such as by air plasma spraying, on the radially outerside 68 b, thereby attaching to the surface of the segment 64.Alternatively, the abradable layer 84 may be pre-manufactured and thensubsequently attached to the radially outer side 68 b in a thermalbonding step. For a repair, a prior abradable layer may be fully orpartially removed and then a new abradable layer or portion may beattached.

Following the attachment of the abradable layer 84 to the segment 64,segment 64 is then mated with the carrier 66 by sliding the segment 74onto the carrier 66. For instance, the dovetail slot 72 of the segment64 is aligned with the dovetail 78 of the carrier 66 and then slid ontothe dovetail 78. As the segment 64 slides onto the carrier, the ridge orridges 86 cut into the abradable layer 84 to form the groove or grooves84 a. In this regard, to facilitate cutting, the ridges 86 may include aknife edge 86 a (FIG. 4A). A knife edge is a tapering of the ridge 86,such as from the base of the ridge, to a high-radius point (tip) thatfocuses the force from the ridge 86 on the abradable layer 84. As willbe appreciated, the length-wise direction of each ridge 86 will besubstantially aligned with the sliding direction so that the ridge 86cuts a uniform groove 84 a, whereas in contrast if the ridge were angledit may scrape across the abradable layer 84 rather than form afunctional groove. In one further example, the groove 84 a may be atleast partially pre-formed prior to contact with the ridge 86. Forinstance, the groove 84 a may be cut or machined into the abradablelayer 84 before sliding the segment 64 onto the carrier 66.

Although a combination of features is shown in the illustrated examples,not all of them need to be combined to realize the benefits of variousembodiments of this disclosure. In other words, a system designedaccording to an embodiment of this disclosure will not necessarilyinclude all of the features shown in any one of the Figures or all ofthe portions schematically shown in the Figures. Moreover, selectedfeatures of one example embodiment may be combined with selectedfeatures of other example embodiments.

The preceding description is exemplary rather than limiting in nature.Variations and modifications to the disclosed examples may becomeapparent to those skilled in the art that do not necessarily depart fromthis disclosure. The scope of legal protection given to this disclosurecan only be determined by studying the following claims.

1. A seal system for a gas turbine engine comprising: a ceramic matrixcomposite (CMC) seal arc segment defining radially inner and outer sidesand having an abradable layer disposed on the radially outer side; acarrier supporting the CMC seal arc segment; a cooling cavity radiallybetween the carrier and the abradable layer; and the carrier including aridge projecting into a groove in the abradable layer and therebyproviding a labyrinth seal that partitions the cooling cavity intosub-cavities, the ridge having a knife edge.
 2. (canceled)
 3. The sealsystem as recited in claim 1, wherein the abradable layer is selectedfrom the group consisting of pure silicon, yttria stabilized zirconia,gadolinia stabilized zirconia, and combinations thereof.
 4. The sealsystem as recited in claim 1, wherein the carrier includes radiallyinner and outer carrier sides, the radially inner carrier side boundsthe sub-cavities, the carrier includes a first supply passage that opensat the radially outer carrier side and at the radially inner carrierside to one of the sub-cavities, and the carrier includes a secondsupply passage that opens at the radially outer carrier side and at theradially inner carrier side to another, different one of thesub-cavities.
 5. The seal system as recited in claim 4, wherein theouter side of the CMC arc segment has a pair of hooks that define adovetail slot there between, the carrier has a dovetail disposed in thedovetail slot, and the first supply passage and the second supplypassage are through the dovetail.
 6. The seal system as recited in claim5, wherein the CMC seal arc segment defines forward and aft sides, oneof the hooks of the pair of hooks is elongated along the forward side,and the other of the hooks of the pair of hooks is elongated along theaft side.
 7. The seal system as recited in claim 1, wherein thesub-cavities are pressurized at different pressures.
 8. The seal systemas recited in claim 1, wherein the carrier, including the ridge, isselected from the group consisting of a Ni-based alloy and acobalt-based alloy.
 9. A gas turbine engine comprising: a compressorsection; a combustor in fluid communication with the compressor section;and a turbine section in fluid communication with the combustor, theturbine section having a seal system disposed about a central axis ofthe gas turbine engine, the seal system including: a ceramic matrixcomposite (CMC) seal arc segment defining radially inner and outer sidesand having an abradable layer disposed on the radially outer side; and acarrier supporting the CMC seal arc segment; a cooling cavity radiallybetween the carrier and the abradable layer; and the carrier including aridge projecting into a groove in the abradable layer and therebyproviding a labyrinth seal that partitions the cooling cavity intosub-cavities, the ridge having a knife edge
 10. The gas turbine engineas recited in claim 9, wherein the carrier includes radially inner andouter carrier sides, the radially inner carrier side bounds thesub-cavities, the carrier includes a first supply passage that opens atthe radially outer carrier side and at the radially inner carrier sideto one of the sub-cavities, the carrier includes a second supply passagethat opens at the radially outer carrier side and at the radially innercarrier side to another, different one of the sub-cavities, the outerside of the CMC arc segment has a pair of hooks that define a dovetailslot there between, the carrier has a dovetail disposed in the dovetailslot, and the first supply passage and the second supply passage arethrough the dovetail.
 11. (canceled)
 12. The gas turbine engine asrecited in claim 9, wherein the sub-cavities are pressurized atdifferent pressures.
 13. The gas turbine engine as recited in claim 12,wherein the abradable layer is selected from the group consisting ofpure silicon, yttria stabilized zirconia, gadolinia stabilized zirconia,and combinations thereof, and the carrier, including the ridge, isselected from the group consisting of a Ni-based alloy and acobalt-based alloy.
 14. (canceled)
 15. (canceled)
 16. (canceled) 17.(canceled)
 18. (canceled)